Ultra compact combustor

ABSTRACT

Embodiments of a combustor for a gas turbine engine are provided herein. In some embodiments, a combustion chamber for a gas turbine engine comprising may include a combustor having an inner volume defined at least partially by a front wall, wherein the wall comprises a plurality of facets each having a through hole fluidly coupled to the inner volume, and wherein the plurality of facets are oriented such that an axis of each of the plurality of facets is offset from a central axis of the combustor by an angle.

CROSS REFERENCE TO RELATED APPLICATION

This patent application claims the benefit of priority, under 35 U.S.C.§ 119, of U.S. Provisional Patent Application Ser. No. 61/989,855, filedMay 7, 2014, titled “ULTRA COMPACT COMBUSTOR” the entire disclosure ofwhich is incorporated herein by reference.

BACKGROUND

The subject matter disclosed herein generally relates to gas turbineengines, and more specifically, to components of combustors.

Conventional gas turbine engines typically include a number ofcomponents configured to direct a flow of air and/or fuel in a desireddirection to facilitate operation of the gas turbine engine. Forexample, as the air and/or fuel flows from one section of the gasturbine engine to the next, the orientation of the flow path may bechanged (“turned”) via or more guide vanes, nozzles, or the like.However, the inventors have observed that redirecting the air/fuel insuch a manner introduces inefficiencies into the operation of the gasturbine engine. Moreover the inclusion of the aforementioned componentsadds weight, cost, and complexity to the gas turbine engine.

Therefore, the inventors have provided an improved gas turbine engine.

SUMMARY

Embodiments of a combustor for a gas turbine engine are provided herein.In some embodiments, a combustion chamber for a gas turbine enginecomprising may include a combustor having an inner volume defined atleast partially by a front wall, wherein the wall comprises a pluralityof facets each having a through hole fluidly coupled to the innervolume, and wherein the plurality of facets are oriented such that anaxis of each of the plurality of facets is offset from a central axis ofthe combustor by an angle.

In some embodiments, a gas turbine engine may include a compressorhaving an exit end; a diffusor disposed downstream of the exit end ofthe compressor; and a combustor disposed downstream of the diffusor, thecombustor having an inner volume defined at least partially by a wall,wherein the wall comprises a plurality of facets each having a throughhole fluidly coupled to the inner volume, and wherein the plurality offacets are oriented such that an axis of each of the plurality of facetsis offset from a central axis of the combustor by an angle.

In some embodiments, a combustion chamber for a gas turbine engine mayinclude a compressor having an exit end; a combustor having an innervolume defined at least partially by a wall, wherein the wall comprisesa plurality of facets each having a through hole fluidly coupled to theinner volume, and wherein the plurality of facets are oriented such thatan axis of each of the plurality of facets is offset from a central axisof the combustor by an angle that is substantially similar to an angleof air flow provided by the exit end of the compressor.

The foregoing and other features of embodiments of the present inventionwill be further understood with reference to the drawings and detaileddescription.

DESCRIPTION OF THE FIGURES

Embodiments of the present invention, briefly summarized above anddiscussed in greater detail below, can be understood by reference to theillustrative embodiments of the invention depicted in the appendeddrawings. It is to be noted, however, that the appended drawingsillustrate only typical embodiments of the invention and are thereforenot to be considered limiting in scope, for the invention may admit toother equally effective embodiments.

FIG. 1 is a cross sectional view of a portion of a conventional gasturbine engine combustion system.

FIG. 1A is a cross sectional view of a portion of the combustion systemshown in FIG. 1

FIG. 2 is a cross sectional view of a portion of the combustion systemshown in FIG. 1 in accordance with some embodiments of the presentinvention.

FIG. 3 is a cross sectional view of a portion of the combustion systemshown in FIG. 1 in accordance with some embodiments of the presentinvention.

FIG. 4 is a cross sectional view of a portion of a combustion system inaccordance with some embodiments of the present invention.

FIG. 4A is a cross sectional view of a portion of the combustion systemshown in FIG. 4 in accordance with some embodiments of the presentinvention.

FIGS. 4B-G are cross sectional views of a portion of the combustionsystem shown in FIG. 4 in accordance with some embodiments of thepresent invention.

FIG. 4H is a cross sectional view of a portion of the combustion systemshown in FIG. 4 in accordance with some embodiments of the presentinvention.

FIG. 5 is a cross sectional view of a conventional gas turbine enginecombustion system.

FIG. 5A is a cross sectional view of the combustion system shown in FIG.5 in accordance with some embodiments of the present invention.

FIG. 6 is a cross sectional view of a portion of the combustion systemshown in FIG. 5 in accordance with some embodiments of the presentinvention.

FIG. 7 is a cross sectional view of a portion of the combustion systemshown in FIG. 5A in accordance with some embodiments of the presentinvention.

To facilitate understanding, identical reference numbers have been used,where possible, to designate identical elements that are common to thefigures. The figures are not drawn to scale and may be simplified forclarity. It is contemplated that elements and features of one embodimentmay be beneficially incorporated in other embodiments without furtherrecitation.

DETAILED DESCRIPTION

Embodiments of a gas turbine engine are disclosed herein. In at leastsome embodiments, the inventive gas turbine engine may reduce oreliminate one or more components typically utilized to direct a flow ofair and/or fuel in a desired direction to facilitate operation of thegas turbine engine, thereby reducing cost, weight and complexity of thegas turbine engine. In addition, in at least one embodiment, theinventive gas turbine engine may include a vortex cavity thatadvantageously provides a helical flow of a fuel/air mixture throughoutthe vortex cavity and combustion chamber, thereby providing an improvedand more efficient mixing and ignition of the fuel/air mixture and,thus, increasing the efficiency of the gas turbine engine. While notintending to be limiting, the inventors have observed that the inventivecomponents of the gas turbine engine disclosed herein may beparticularly suitable for use in combustors.

Referring to FIG. 1, in some embodiments, a conventional gas turbineengine combustion system (combustion system)100 generally includes acombustion chamber 102 having an inner liner 144 and an outer liner 104that at least partially defines a combustor 108. In some embodiments acasing 140 may be disposed about the outer liner 104. When present thecasing 140 and an outer portion 136 of the outer liner 104 may form anouter passage 138. Although only a portion is shown, it is to beunderstood that the combustion chamber 102 may be annular in form, forexample about an axis 110.

In some embodiments, air is directed into the combustion system 100 viaan intake that includes, for example, one or more fans, compressors, orthe like (e.g., partial view of compressor rotor 112 shown). The airflows from the intake, via an exit end 114 of the compressor and througha diffuser 106. The diffusor 106 is configured to direct the air towardsthe combustion chamber 108. In some embodiments, one or more guide vanesor struts (e.g., such as collectively shown at 116) may be disposedwithin the diffusor 106 and/or proximate the exit end 114 of thecompressor.

In some embodiments, an inlet module 120 comprising a cowl 126 may bedisposed about a through hole 134 formed in a front wall 142 of thecombustion chamber 108 and configured to direct the air from thediffusor 106 to an inner volume of the combustion chamber 108. Althoughonly one inlet module 120, cowl 126 and through hole 134 are shown inthe figure, multiples of each may be present.

In some embodiments, one or more swirlers 128 may be disposed proximatethe through hole 134 to facilitate mixing of the air and a fuel providedby a fuel source 122 via a fuel injector 124 for ignition andcombustion. In some embodiments, a splash plate 130 may be disposedwithin the combustion chamber 108 to reduce instances of damage toportions of the front wall 142 caused by increased temperaturesresulting from ignition of the fuel/air mixture.

In some embodiments, following ignition, the air is directed out of thecombustion chamber 108 and towards one or more turbines (e.g., partialview of one turbine rotor 118 shown) via a turbine nozzle 132 (stage onenozzle). When present, the turbine nozzle 132 functions to direct a flowof the air at a desired angle in a desired flow path. In someembodiments, the turbine nozzle 132 may include a plurality of throughholes that allows a flow of cooling air to prevent heat induced damage(e.g., deformation, melting, or the like) to the turbine nozzle 132.Although only one turbine is shown, it is to be understood that morethan one may be present, for example such as one or more low pressureturbines, high pressure turbines, or the like.

The inventors have observed that conventional combustors typicallyinclude a number of mechanisms to direct air flow between components tofacilitate an efficient use of such air. For example, referring to FIG.1A the diffusor 106 may include at least one of outlet guide vanes 146,aerodynamically shaped struts 148, or the like, to alter an angle of theair flow, reduce or eliminate a swirling component of the air flowand/or direct a desired portion of the air towards the inlet module 120of the combustion chamber 108 (exemplary air flow shown in phantom).However, the inventors have observed that directing the air flow in sucha manner results in a high pressure drop and losses in air flow, therebyreducing efficiency of the gas turbine engine (e.g., reduced Braytoncycle efficiency). Moreover, inclusion of the components needed tofacilitate directing the air flow (e.g., the one or more guide vanes)adds cost and complexity to the combustion chamber 102.

As such, referring to FIG. 2, in some embodiments the diffusor 106 maybe configured having an inner surface 206 and/or struts 148 angled suchthat the flow of air flowing through the diffusor 106 flows in a desiredflow path. In some embodiments, the angle 204 may be substantiallysimilar to a swirl angle of a flow of air as it exits the exit end 114of the compressor 112. For example, in some embodiments, the diffusor106 may be configured such that a central axis 208 of the diffusor isoffset from the central axis 110 of the combustion chamber 102 by anangle 204, thereby providing a flow path 202 that is parallel to acentral axis 208 of the diffusor 106. The angle 204 may be any anglesuitable to direct the flow of air in a desired flow path, for example,such as about 15 degrees to about 60 degrees, or in some embodiments,about 15 degrees to about 45 degrees.

By configuring the diffusor 106 as shown in FIG. 2 the inventors haveobserved that components that would otherwise be present in the diffusor(e.g., the outlet guide vanes 146) may be eliminated, thereby reducingcomplexity and length of the diffusor. In addition, by eliminating theoutlet guide vanes 146 the swirling component of the airflow may bemaintained, thereby decreases losses in airflow that otherwise occurswhen removing the swirling component in conventional turbine engines.Moreover, the pressure drop typically observed in conventionalcombustors may be reduced, thereby also decreasing losses in the airflow and increasing the efficiency of the gas turbine engine.

The inventors have observed that conventional combustion chamberstypically include one or more through holes and inlet modules disposedon a singular surface (front wall) of the combustion chamber, forexample such as described above with respect to FIG. 1. However, becauseof the orientation of the front wall of the combustion chamber, the flowpath of the air typically needs to be altered to facilitate a desiredflow of air into the combustion chamber.

As such, referring to FIG. 3, in some embodiments, the front wall 142 ofthe combustor 108 may comprise a plurality of faces (facets) 302 thatare positioned or oriented to receive the flow of air at a desiredangle. The plurality of faces 302 may be positioned in any manner toreceive the air flow such as the corrugated or step-like configurationshown in FIG. 3. For example, in some embodiments, each of the pluralityof faces 302 may be oriented such that an axis 324 of each of theplurality of faces (facets) 302 is offset from a central axis of thecombustor by a desired angle (e.g., angle 204). The desired angle may beany angle, for example, such as about 15 degrees to about 60 degrees, orin some embodiments, about 15 degrees to about 45 degrees. For example,in some embodiments, each of the faces 302 may be positioned such thatthe faces 302 are substantially perpendicular to the air flow 202provided by the diffusor 106. In such embodiments, the angle 204 may besubstantially similar to a swirl angle of a flow of air as it exits theexit end 114 of the compressor 112.

The inventors have observed that by providing the plurality of faces302, the flow of air follows the flow path 202 dictated by the diffusor(described in FIG. 2) thereby further reducing the need to alter thedirection of the flow of air, thus eliminating the need for additionalcomponents (e.g., outlet guide vanes, etc.) that would otherwise bepresent in conventional configurations.

In some embodiments, each face or facet 302 may comprise a through hole310, one or more air swirlers 308 and one or more guide vanes 306. Insome embodiments, a splash plate (e.g., such as the splash plate 142shown in FIG. 1) may be disposed on an interior surface 314 of each ofthe faces 302. In some embodiments, fuel may be provided proximate eachface 302 via a fuel injector 316. In operation, the air and fuel mixturemay be ignited via an igniter (not shown) proximate the splash plate 312and subsequently flow into the combustion chamber (ignition andsubsequent flow indicated at 320 and 318 respectively).

Referring to FIG. 4, in some embodiments, the combustion chamber 102 maycomprise a vortex cavity (cavity) 402 fluidly coupled to the combustor108. When present, the vortex cavity 402 may function to provide adesired air/fuel mixture in a desired distribution and to stabilize aflame formed by the ignited air/fuel mixture. Although only one vortexcavity 402 is shown in the figure, any number of vortex cavitiessuitable to provide a desired air/fuel mixture in a desired distributionand to stabilize a flame formed by the ignited air/fuel mixture may bepresent.

Although shown as a singular front wall 142 in FIG. 4, the front wall142 may be configured to have a plurality of faces or facets, such asdescribed above with respect to FIG. 3. Alternatively, in someembodiments, the guide vanes 306 may be configured such that a centralaxis 430 of the guide vanes 306 may be offset from the front wall 142 byan angle 434, for example, such as shown in FIG. 4A. The angle 434 maybe any angle sufficient to accommodate for a flow of air directedtowards the combustor (e.g., from the diffusor discussed above).

Referring to FIG. 4, the vortex cavity 402 generally comprises aplurality (e.g., three) of sides 432 and an open end 414 that is influid communication with the combustor 108. Although the three sides 432are shown as connected at an angle, in some embodiments, one or moreinner surfaces 436 between the sides 432 may be rounded. When present,the rounded inner surfaces may function to reduce mechanical or thermalstresses within the vortex cavity 402 and/or may eliminate areas havinga trapped flow (e.g., eddy flow) or no flow.

In some embodiments, a plurality of through holes (injection holes) 404(four shown) may be formed in at least one of the sides 432 of thevortex cavity 402. When present, the plurality of injection holes 404directs a flow of air into vortex cavity 402 to facilitate the formationof the vortex 412. The injection holes 404 may be configured in anymanner suitable to facilitate, for example, the creation of the vortex412, a desired vortex shape, fuel/air mixture, movement of the vortexflow into the combustion chamber 108, or the like. In addition, in anyof the embodiments described herein, the injection holes 404 may includea mechanism (shown in phantom at 450), for example, such as a nozzle orthe like to facilitate varying one or more flow characteristics (flowrate, pressure, direction or the like) through the injection holes 404.Moreover, although shown in the cross sectional views as at leastsubstantially circular, it is to be understood that the injection holes404 may have any shape, for example such as slotted holes or the like,suitable to provide a desired air flow.

Referring to FIG. 4B, in some embodiments, the vortex cavity 402 mayinclude one or more protrusions (three protrusions 416 shown) extendinginto the vortex cavity 402 to direct the flow of air about the vortexcavity 402 in a circular flow path to further facilitate the formationof the vortex 412, for example, such as shown in the figure. In someembodiments, the vortex cavity 402 may further comprise a fuel injectionassembly 418 configured to facilitate the insertion of a fuel injector410 into the vortex cavity 402. In some embodiments, the fuel injectionassembly 418 may generally comprise a collar 422 coupled to an outersurface 424 of the vortex cavity 402 and a guide 420 disposed within thecollar 422.

Referring back to FIG. 4, a fuel source 408 provides fuel to the vortexcavity 402 via the fuel injector 410 disposed within the vortex cavity402. The fuel injector 410 provides the fuel from the fuel source 408 tothe via a fuel injector hole 452. In some embodiments, the fuel injector410 may comprise a mechanism (shown in phantom at 462 of FIG. 4H)disposed in or about the fuel injector hole, for example, such as anozzle or the like to facilitate providing the fuel and/or varying oneor more flow characteristics (flow rate, pressure, direction or thelike) of the fuel. In some embodiments, the fuel injector 410 isdisposed within the vortex cavity 402 such that an axis 464 of the fuelinjector hole 452 is parallel to a tangential component 466 of thevortex cavity 402, for example, such as shown in FIG. 4H. Disposing thefuel injector 410 in such a manner allows for fuel injector 410 toprovide the fuel in a flow path that is generally tangential (e.g., in adirection of the axis 464 of the fuel injector hole 452) to the vortexcavity 402. The inventors have observed that providing the fuel in sucha manner may at least partially facilitate a movement of the fuel/airmoisture in a helical path through the vortex cavity 402.

In some embodiments, configurations of the injections holes 404 mayinclude varied placement, number or directionality/angle of eachinjection hole with respect to the vortex cavity 402. In addition, aflow rate through each of the injection holes 404 may be variedindependently. The inventors have observed that varying such flow rate,placement, number or directionality of the injection holes may provide amechanism to facilitate formation of the vortex and/or forming thevortex or flow of air having desired characteristics. For example,directing flow towards the combustion chamber 108 or increasing a flowrate of the air towards the combustion chamber 108 may facilitatedirecting the vortex flow to the combustion chamber 108 (vortextransport), thereby moving a point of combustion towards the combustionchamber 108. In another example, directing flow towards a top or closedend of the vortex cavity 402 or parallel with the top of the vortexcavity 402 may facilitate the formation or increase a rotationalcomponent of the vortex 412, thereby facilitating a desired mixing ofthe fuel and air.

In addition to the above, in some embodiments, one or more of thethrough holes 404 may be configured such that a flow path of air 426provided by the through holes 404 may have a tangential component 470and a radial component 468. As used herein, “radial” may refer todirections that are radially inward or outward with respect to a centerof an annular shape of the cavity 402 or combustor 108 and “tangential”may refer to a tangential direction at any point about of the annularshape of the cavity 402 or combustor 108. The inventor have observedthat providing the air via the through holes 404 in such a manner mayfunction to reduce or eliminate a need to further turn the air flow asit exits the combustion chamber 108. Eliminating the need to furtherturn the air flow allows for a reduction of a length, or in someinstances, elimination, of a turbine nozzle (first stage nozzle) thatwould otherwise be required to turn the air flow.

An exemplary illustration of an exemplary helical path resulting fromthe flow path 426 of air provided by the plurality of injection holes404 or the plurality of injection holes 404 in combination of the flowpath 428 of fuel provided by the fuel injector 410 is shown in FIG. 4H.The inventors have observed that the helical flow of air/fuel mayadvantageously provide an improved and more efficient mixing andignition of the fuel/air mixture and, thus, increasing the efficiency ofthe gas turbine engine. Moreover, the inventors have observed that thehelical flow of air/fuel may advantageously providing a desireddistribution of the fuel/air mixture throughout the combustion chamber108, thereby further increasing the efficiency of the gas turbineengine. In addition, the inventors have observed that the helical flowof air/fuel may advantageously accommodate for a pressure gradientformed within the combustion chamber 108 (e.g., caused by an area ofhigh pressure proximate the outer liner 104 and an area of low pressureproximate the inner liner 144), thereby further providing theaforementioned advantages.

FIGS. 4 and 4B-G depict exemplary illustrative embodiments of variousconfigurations of the injection holes 404. Although a number ofinjection holes are depicted in the cross sectional views of the vortexcavity of FIGS. 4 and 4A-G, it is to be understood that injection holesmay be provided throughout the vortex cavity 402, for example such asshown in FIG. 4H. In addition, although shown as separateconfigurations, multiple configurations, or combination of theillustrated configurations may be utilized throughout the vortex cavity402.

In one example, in some embodiments, the vortex cavity 402 may include afirst side (aft) 444 of the vortex cavity 402 having a plurality ofinjection holes 404 (e.g., injection hole 454 and injection hole 456), asecond side (top or closed end) 446 having an injection hole 458 and athird side (fore) 448 having an injection hole 460, for example, such asshown in FIG. 4. In such embodiments, each of the injection holes 404may be configured to provide a desired directionality of air flow tofacilitate, for example, formation of the vortex 412, movement of thevortex 412, or the like. For example, a first injection hole (injectionhole 454) of the first side 444 may be configured such that an flow isdirected parallel to the top 446 of the vortex cavity 402 and a secondinjection hole (injection hole 456) of the first side 444 may beconfigured such that an air flow is directed generally towards the top446 of the vortex cavity 402. In addition, the injection hole 458 of thetop 446 may be configured such that an air flow is directed generallytowards the combustion chamber 108 and the injection hole 460 of thethird side 448 may be configured such that an air flow is directedparallel to the top 446 of the vortex cavity 402. FIG. 4C depicts anexemplary configuration of the vortex cavity 402 that is similar to theconfiguration shown in FIG. 4, however, without the first injection hole(injection hole 454) of the first side 444 of the vortex cavity 402.FIG. 4D depicts an exemplary configuration of the vortex cavity 402 thatis similar to the configuration shown in FIG. 4, however, without thesecond injection hole (injection hole 456) of the first side 444. FIG.4E depicts an exemplary configuration of the vortex cavity 402 that issimilar to the configuration shown in FIG. 4C, however, having theinjection hole 460 of the third side 448 configured such that an airflow is directed towards the combustion chamber 108. FIG. 4F depicts anexemplary configuration of the vortex cavity 402 that is similar to theconfiguration shown in FIG. 4C, however, having the injection hole 460of the third side 448 configured such that an air flow is directedtowards the top 446 of the vortex cavity 402.

Referring to FIG. 4G, in some embodiments, an angle 442 between acentral axis 438 of the vortex cavity 402 and a central axis 440 of thecombustion chamber 108 may be varied to facilitate altering a swirlangle of the air/fuel mixture through the combustion chamber 108. Theinventors have observed that altering the swirl angle of the air/fuelmixture through the combustion chamber 108 may eliminate the need tofurther turn the air flow as it exits the combustion chamber 108.Eliminating the need to further turn the air flow allows for a reductionof a length, or in some instances, elimination, of a turbine nozzle(first stage nozzle) that would otherwise be required to turn the airflow.

As discussed above, the inventors have observed that conventionalturbine engines typically require multiple components (e.g., one or moreguide vanes, diffusors, or the like) to change an orientation of a flowpath of air to facilitate a desired flow path through the combustor. Forexample, referring to FIG. 5, a conventional gas turbine engine 500 mayinclude an axial compressor 504, a centrifugal compressor 506, adiffusor 508, and an inner liner 512 and outer liner 502 that at leastpartially defines a combustor 510. In some embodiments, a casing 514 maybe disposed about the outer liner 502. When present, the casing 514 andan outer portion 516 of the outer liner 502 may form an outer passage518.

In operation, the axial compressor 504 receives air from an intake (notshown). The air is compressed and is received by the centrifugalcompressor 506. The diffusor 508 receives the compressed air from thecentrifugal compressor 506 and directs a desired portion of thecompressed air to the combustor 510. In some variations the diffusor 508may include one or more guide vanes (one guide vane assembly 520 shown)that functions to alter an angle of the air flow, reduce or eliminate aswirling component of the air flow and/or direct a desired portion ofthe air towards the combustor 510. The compressed air is mixed with afuel and ignited within the combustor 510. Following ignition, the airis directed out of the combustor 510 and towards one or more turbines(e.g., partial view of one turbine rotor 526 shown) via a turbine nozzle522 (stage one nozzle).

The inventors have observed that providing the air radially (e.g., viathe centrifugal compressor 506 and diffusor 508 as shown in FIG. 5)reduces the need for the orientation of the air flow to be altered(turned) prior to exiting the combustor, thereby increasing theefficiency of the turbine engine. However, the inventors have observedthat such configurations still require one or more components (e.g.,diffusor, guide vanes, turbine nozzle 522, or the like) to alter theflow of air to facilitate operation of the turbine engine.

As such, referring to FIG. 5A in some embodiments, the inner liner 512and outer liner 502 may be curved such that the combustor 510 comprisesa first portion 534 fluidly coupled to a second portion 536, wherein acentral axis 538 of the first portion 534 is offset from the centralaxis 540 of the second portion 536 by an angle 542. Configuring thecombustor 510 in such a manner allows for an orientation and swirlingcomponent of the flow of air provided by the diffusor 508 to bemaintained as it flows through the combustion chamber 506, therebyeliminating the need to further turn the air flow as it exits thecombustion chamber 506. Eliminating the need to further turn the airflow allows for a reduction of a length, or in some instances,elimination, of a turbine nozzle (first stage nozzle) that wouldotherwise be required to turn the air flow prior to reaching the turbine526. Moreover, configuring the combustor 510 such as shown in FIG. 5Areduces an overall length of the combustor 510, thereby providing aturbine engine having a reduced footprint. Reducing the size and/oreliminating the aforementioned components advantageously reduce theoverall weight, cost and complexity of the turbine engine.

The first portion 518 and second portion 520 of the combustor 506 may bedisposed in any position with respect to one another suitable tofacilitate the aforementioned desired air flow. For example, in someembodiments, the central axis 514 of the first portion 518 may be offsetfrom the central axis 516 of the second portion 520 by about 45 to about180 degrees, or in some embodiments, about 90 degrees, such as shown inFIG. 5A.

In some embodiments, a cavity 524 may be fluidly coupled to thecombustor 510. In some embodiments, the cavity 524 may be disposeddirectly in-line with the combustor 510 (such as indicated in phantom at544) or disposed such that at least a portion of the cavity 524 isdisposed on a side of the combustor 510 (such as indicated at 524). Whenpresent, the cavity 524 may be similar in configuration and function tothe vortex cavity 402 described above. In some embodiments, thecombustor 510 may include one or more guide vanes 528 configured todirect a desired flow of the compressed air provided by the diffusor 508into the combustor 510 via one or more through holes 532 formed in afront wall 530 of the combustor 510 (e.g., similar to the one or moreguide vanes 306 described above).

Referring to FIG. 6, in some embodiments, the front wall 530 may becurved to accommodate for a direction of air flow are various positionsacross the front wall 530 (e.g., air flow shown by arrows 602, 604,606). In such embodiments, each of the guide vanes 528 may be angled tofurther accommodate for the direction of air flow. In some embodiments,the curvature of the front wall 530 and/or the orientation of each ofthe guide vanes 528 may be dependent on a curvature of the terminal end608 of the diffusor 508 to facilitate an efficient capture of airprovided by the diffusor 508.

Alternatively, or in combination, in some embodiments, each guide vane706 may be movable with respect to the front wall 530, for example suchas shown in FIG. 7. In such embodiments, the guide vane 706 may becoupled to a spherical or cylindrical base 702 having a through hole 704formed there through. The base 702 may be disposed within a collar 708that is coupled to the front wall 530, thereby allowing the sphericalbase, and therefore the guide vane 706, to move in a direction and to adegree of freedom dictated by the collar 708 (movement shown by arrow710). In some embodiments, the guide vane 706 may be moved mechanically,for example via a conventional actuator ring disposed about the turbineengine. Providing a movable guide vane 706 allows the guide vane 706 toaccommodate for a changing direction in air flow, thereby providingadditional degrees of freedom and an increased window of operability.

Thus, embodiments of a gas turbine engine have been provided herein. Inat least some embodiments, the inventive gas turbine engine may reduceor eliminate one or more components typically utilized to direct a flowof air and/or fuel in a desired direction to facilitate operation of thegas turbine engine, thereby reducing cost, weight and complexity of thegas turbine engine. In addition, in at least some embodiments, the gasturbine engine may include an inventive vortex cavity that facilitates ahelical flow of a fuel/air mixture that advantageously provides animproved and more efficient mixing and ignition of the fuel/air mixtureand, thus, increasing the efficiency of the gas turbine engine.

Ranges disclosed herein are inclusive and combinable (e.g., ranges of“about 0 psi to about 25,000 psi”, is inclusive of the endpoints and allintermediate values of the ranges of “about 0 psi to about 25,000 psi,”etc.). “Combination” is inclusive of blends, mixtures, alloys, reactionproducts, and the like. Furthermore, the terms “first,” “second,” andthe like, herein do not denote any order, quantity, or importance, butrather are used to distinguish one element from another, and the terms“a” and “an” herein do not denote a limitation of quantity, but ratherdenote the presence of at least one of the referenced item. The modifier“about” used in connection with a quantity is inclusive of the statevalue and has the meaning dictated by context, (e.g., includes thedegree of error associated with measurement of the particular quantity).The suffix “(s)” as used herein is intended to include both the singularand the plural of the term that it modifies, thereby including one ormore of that term (e.g., the colorant(s) includes one or morecolorants). Reference throughout the specification to “one embodiment”,“some embodiments”, “another embodiment”, “an embodiment”, and so forth,means that a particular element (e.g., feature, structure, and/orcharacteristic) described in connection with the embodiment is includedin at least one embodiment described herein, and may or may not bepresent in other embodiments. In addition, it is to be understood thatthe described elements may be combined in any suitable manner in thevarious embodiments.

While the invention has been described with reference to exemplaryembodiments, it will be understood by those skilled in the art thatvarious changes may be made and equivalents may be substituted forelements thereof without departing from the scope of the invention. Inaddition, many modifications may be made to adapt a particular situationor material to the teachings of the invention without departing fromessential scope thereof. Therefore, it is intended that the inventionnot be limited to the particular embodiment disclosed as the best modecontemplated for carrying out this invention, but that the inventionwill include all embodiments falling within the scope of the appendedclaims.

1.-20. (canceled)
 21. A combustion chamber for a gas turbine enginecomprising: a compressor having an exit end; a combustor having an innervolume defined at least partially by a wall, wherein the wall comprisesa plurality of facets each having a through hole fluidly coupled to theinner volume, and wherein the plurality of facets are oriented such thatan axis of each of the plurality of facets is offset from a central axisof the combustor by an angle that is substantially similar to an angleof air flow provided by the exit end of the compressor, the combustorfurther comprising an air swirler disposed about each of the throughholes; and an inlet guide vane disposed about each of the through holesand each air swirler.
 22. A combustor for a gas turbine enginecomprising: a combustion chamber having an inner volume defined at leastpartially by a front wall, wherein the front wall comprises a pluralityof cylindrical or spherical bases each having a through hole fluidlycoupled to the inner volume, and wherein each of the plurality ofcylindrical or spherical bases is disposed within a collar that iscoupled to the front wall, wherein each cylindrical or spherical base ismovable within each collar through at least one angle; and an inletguide vane disposed through each of the through holes of the cylindricalor spherical bases.